This invention relates generally to gas turbine engines and, more particularly, to methods and systems for cooling integral turbine shroud assemblies.
One known approach to increase the efficiency of gas turbine engines requires raising the turbine operating temperature. However, as operating temperatures are increased, the thermal limits of certain engine components may be exceeded, resulting in reduced service life and/or material failure. Moreover, the increased thermal expansion and contraction of components may adversely affect component clearances and/or component interfitting relationships. Consequently, cooling systems have been incorporated into gas turbine engines to facilitate cooling such components to avoid potentially damaging consequences when exposed to elevated operating temperatures.
It is known to extract, from the main airstream, air from the compressor for cooling purposes. To facilitate maintaining engine operating efficiency, the volume of cooling air extracted is typically limited to only a small percentage of the total main airstream. However, this requires that the cooling air be utilized with the utmost efficiency in order to maintain the temperatures of components within safe limits.
For example, one component that is subjected to high temperatures is the shroud assembly located immediately downstream of the high pressure turbine nozzle extending from the combustor. The shroud assembly extends circumferentially about the rotor of the high pressure turbine and thus defines the outer boundary (flow path) of the main gas stream flowing through the high pressure turbine. Gas turbine engine efficiency is negatively affected by a fluctuation in turbine blade clearance measured between a radially outer surface of the turbine blade and a radially inner surface of the shroud assembly. During transient engine operation, the turbine blade clearance is a function of the relative radial displacements of the turbine rotor blade and the shroud assembly. The turbine rotor blade typically has a larger mass than the stationary shroud system and, thus, during turbine operation, the turbine rotor blade typically has a slower thermal response than the shroud assembly. When the difference in the turbine rotor radial displacement and the shroud assembly radial displacement is too great, the blade clearance is increased, which results in a reduction in engine efficiency.